This invention relates to coolable airfoils of the type used in high temperature rotary machines such as gas turbines and, more particularly, to an improved cooling scheme for airfoils with internal cooling passages with enhanced efficiency.
The turbine of a gas turbine engine or machine is subjected to extremely high temperatures. The temperature of hot gases entering the turbine from the combustor is generally well above the melting point temperatures of the alloys from which turbine rotor blades and stator vanes are fabricated. Since both blades and vanes are subjected to such high temperatures, they must be cooled to maintain their structural integrity.
Generally, blades and vanes are cooled by air bled from the engine""s compressor, bypassing the combustor. The cooling air then flows through internal cavities of the respective blades and vanes. The air temperature of the air bled from the compressor is generally at a relatively lower temperature than the temperature of the hot gases. It will be understood that compressor bleed air for such cooling will be unavailable to support combustion in the combustor. Thus, to minimize any sacrifice in engine performance due to inadequate airflow to support combustion, any scheme for cooling blades and vanes must optimize the utilization of compressor bleed cooling air. Airfoil cooling is accomplished by external film cooling and internal air impingement and convection cooling, or a combination of both.
In convection cooling, compressor bleed air flows through the internal cavities of the blades and vanes, continuously removing heat therefrom. Compressor bleed air enters the cavities through one or more inlets which discharges into the internal cavities. The internal cavities may include fins or ridges (also known as xe2x80x9ctrip stripsxe2x80x9d) in a wall thereof, which facilitate improved convection cooling of the walls of the blades and vanes.
Film cooling has been shown to be very effective but requires a great deal of fluid flow. Also, the fabrication and machining of an airfoil with film cooling holes adds a degree of complexity that is costly.
It will also be appreciated that once the cooling air exits the internal cavity of the airfoil and mixes with the hot gases, a severe performance penalty is incurred due to the mixing process and the different temperature levels of the mixing flows. This undesirable effect is heightened if the cooling air is ejected from film holes located on the suction side of the airfoil generally in the trailing edge region as there are significant mixing losses due to the adverse pressure gradients that are formed in this region. If film cooling holes are placed beyond the throat area of the suction side wall, then undesirable flow separation is also possible. Thus, film cooling requires a greater amount of cooling air than with the possibility of inadequate cooling of the outer surfaces of the airfoil.
If film cooling holes are not employed, the airfoil can creep due to lack of cooling. If the airfoil is coated with a thermal barrier coating, for example, the coating can spall, leaving the metal exposed to the hot gases with the result being undesirable cracking or burning of the airfoil walls. One of the traditional approaches to address the creep, is to increase the amount of heat transfer by using protruding ribs (trip strips, turbulators) in the internal passages of the blades and vanes to promote turbulent mixing in the bulk flow. However, in the trailing edge region, the internal cavities are relatively small thus making it difficult to add protruding ribs without causing flow blockage of the spent cooling flow within the internal cavity.
Impingement cooling is another cooling technique that may be employed to alleviate creep; however, it also has it drawbacks. With air impingement, compressor bleed air is channeled to the inside of an airfoil and directed onto the inside walls of the airfoil. The air then exits the airfoil through a set of film holes provided within the airfoil walls. However, if impingement cooling is employed in a region where it is not desirable to utilize film cooling, then the spent impingement air does not exit through proximate film cooling holes. This results in the reduction of the impingement cooling effectiveness as the cross flow from upstream impingement holes degrades the impingement action of the downstream impingement holes.
Therefore, there is a need in the art for an airfoil with an optimized cooling scheme which extends airfoil life and in turn, optimizes the efficiency of the engine and also reduces the amount of fuel burned by the engine, thus enhancing the economy of operation of the engine.
The above discussed and other drawbacks and deficiencies are overcome or alleviated by the present invention.
Accordingly, the present invention provides a turbine airfoil having enhanced heat transfer for cooling the internal cavity of an airfoil which optimizes the efficiency of the engine by minimizing the amount of compressor bleed air required.
The inner surface of the aft internal cavity, or any other cavity, of the airfoil is convectively cooled using a plurality of indentations positioned along the inner surface of the airfoil cavity. More particularly, the indentations may be located along the inner surface of the convex (suction) side wall of the airfoil proximate to the trailing edge. In this way, the present invention provides for the cooling of the inner surface of the airfoil cavities which extends the life of the airfoil without requiring a supply of additional cooling air, as would be the case if a film cooling scheme were employed. As a consequence, turbine efficiency is not adversely affected. Advantageously, this cooling scheme also does not block the spent impingement flow from flowing within the airfoil internal cavities. More particularly, the indentations when located in the aft internal cavity proximate to the trailing edge do not prevent the spent impingement flow from flowing out of the aft internal cavity and through the cooling slots of the trailing edge.
It is preferred that the indentations are staggered extending in the spanwise or longitudinal direction of the airfoil. It is most preferred that the indentations are arranged in a staggered array in at least two longitudinally extending rows such that the plurality of indentations are centered along a single zig-zag line. The pattern of longitudinal placement of the indentations could be parallel to the cooling air, perpendicular to such flow, or to any other angle to the cooling flow. Preferably, the pattern of the indentations would be optimized with respect to the direction of the local flow streamlines to provide the highest heat transfer surface enhancement possible.
Preferably, each indentation is a dimple that extends into the inner surface of the airfoil. However, the indentation could have alternative geometric configurations which can produce the same heat transfer enhancement in the trailing edge including all of the additional benefits that will are detailed herein. The indentations may have parallelepiped, elliptical, kidney, or rectangular shapes. Alternatively, the indentations may be ramps, semi-circular or race-track shapes. The selected indentation geometric configuration depends on the area and the desired heat transfer enhancement. Further, the indentations can also be of varying depth that is optimized in relation to the spacing between adjacent indentations.
The present invention also contemplates a method of enhancing the cooling of a turbine airfoil by forming a plurality of indentations on the inner surface of an internal cavity within the airfoil. This method can also be employed in an exiting airfoils to further enhance existing cooling schemes.